Hydrogen powered aircraft

ABSTRACT

Disclosed is an aircraft, configured to have a wide range of flight speeds, consuming low levels of power for an extended period of time, while supporting a communications platform with an unobstructed downward-looking view. The aircraft includes an extendable slat at the leading edge of the wing, and a reflexed trailing edge. The aircraft comprises a flying wing extending laterally between two ends and a center point. The wing is swept and has a relatively constant chord. The aircraft also includes a power module configured to provide power via a fuel cell. The fuel cell stores liquid hydrogen as fuel, but uses gaseous hydrogen in the fuel cell. A fuel tank heater is used to control the boil-rate of the fuel in the fuel tank. The fuel cell compresses ambient air for an oxidizer, and operates with the fuel and oxidizer at pressures below one atmosphere. The aircraft of the invention includes a support structure including a plurality of supports, where the supports form a tetrahedron that affixes to the wing.

BACKGROUND OF THE INVENTION

The present application is a continuation-in-part of application Ser.No. 10/073,828, filed Feb. 11, 2002 now abandoned, is a divisional ofapplication Ser. No. 09/826,424 filed Apr. 3, 2001 now U.S. Pat. No.6,550,717, which claims priority from two U.S. provisional patentapplications, Ser. No. 60/194,137, filed Apr. 3, 2000, and Ser. No.60/241,713 now U.S. Pat. No. 6,550,717, filed Oct. 18, 2000, all four ofwhich are incorporated herein by reference for all purposes.

This invention relates generally to aircraft and their componentsystems, and, more particularly, to improved high-performance aircraftsystems capable of high-altitude station keeping within tight altitudeand perimeter boundaries for extended periods of time.

A worldwide expansion in the demand for communication bandwidth isdriving up the bandwidth requirements between satellites andground-stations. One way to increase this satellite-to-ground bandwidthis to interpose one or more high-altitude platforms (HAPs) configuredfor relaying signals between the two. A HAP allows for lower powertransmissions, narrower beamwidths, as well as a variety of otheradvantages that provide for greater bandwidth. However, due to ademanding set of design requirements, years of design efforts atcreating highly effective HAPs are only now beginning to come tofruition.

In particular, it is desirable to have a stratospheric aircraft, capableof carrying a significant communications payload (e.g., a payload ofmore than 100 kg that consumes more than 1 kw of electric power), thatcan remain aloft for days, weeks or even months at a time. This flightcapability will preferably be maintainable even in zero or minimumsunlight conditions where solar power sources have little functionality.Also, the aircraft is preferably remotely pilotable to limit its weightand maximize its flight duration.

The communications payload preferably is configured to view downwardover a wide, preferably unobstructed field of view. The aircraft willpreferably be capable of relatively high-speed flight that is adequateto travel between its station and remote sites for takeoff and/orlanding to take advantage of benign weather conditions. At the sametime, the aircraft preferably is capable of maintaining a tight,high-altitude station in both high-wind and calm conditions, thusrequiring relatively high-speed and relatively low-speed flight, and asmall turning radius while maintaining the payload's downward-looking(and preferably upward-looking for some embodiments) view. To meet thesestringent design specifications, the performance of the aircraft's powersystem, flight control system and airframe configuration and are allpreferably improved over prior practice.

Power Systems

Conventional aircraft are typically powered using aviation fuel, whichis a petroleum-based fossil fuel. The prior art mentions the potentialuse of liquid hydrogen as a fuel for manned airliners and supersonicstratospheric flight. There is also 25-year-old prior art mentioning thepossibility of using liquid hydrogen as fuel for a stratospheric blimp.

U.S. Pat. No. 5,810,284 (the '284 patent), which is incorporated hereinby reference for all purposes, discloses an unmanned, solar-poweredaircraft that significantly advanced the art in long-duration,stratospheric aircraft. It flies under solar power during the day, andstores up additional solar power in a regenerative fuel cell battery foruse during the night to maintain its station. The fuel cell battery is aclosed system containing the gaseous elements of hydrogen and oxygenthat are dissociated from, and combined into, water.

The aircraft disclosed in the '284 patent is an unswept, span-loaded,flying wing having low weight and an extremely high aspect ratio.Multiple electric engines are spread along the wing, which issectionalized to minimize torsion loads carried between the sections.Most or all of the sections contain a hollow spar that is used tocontain the elements used by the fuel cell. Large fins extend downwardfrom inner ends of the sections. The wings contain two-sided solarpanels within transparent upper and lower surfaces to take maximumadvantage of both direct and reflected light.

The above-described technologies cannot provide for long-duration,high-altitude flights with tight stationkeeping when the available solarpower is highly limited.

Flight-Control Components

Various components are known for use in controlling flight. Eachcomponent has unique advantageous and disadvantageous characteristics.

Many present-day small aircraft and some sailplanes use simple flaps toincrease camber and obtain higher lift coefficients, and hence, adequatelift at lower speeds. Such flaps are typically retracted or faired toreduce drag during high-speed flight, and also during turbulence toreduce the maximum G loads that the wing will then experience. Animportant characteristic of the use of flaps, or of the use of highlycambered airfoils designed for high lift, is that the extended flap orhighly cambered airfoil provides the wing with a large negative pitchingmoment. This affects both overall vehicle stability and the wing'storsional twisting. Indeed, for high aspect-ratio wings, the twist atthe wing's outer portions due to a negative pitching moment can posesevere structural and flight control problems.

Airliners use both leading edge slats and sophisticated flaps, such asslotted or Fowler flaps, to widen their speed range. Small planes employslats that open automatically when needed. Hang gliders have employedflexible airfoil tightening to decrease camber for high-speed flight.Some work has been done with flexible flap material that unrolls andpulls back from the rear of the wing. Some aircraft feature wingscharacterized by a sweep that can be varied in flight, even turning theentire wing so that it is not perpendicular to the flight directionduring high-speed flight.

For maintaining low-speed flight without stalling, large solid or poroussurfaces that hingedly swing up from a wing top in low-speed flight topotentially stabilize vortices immediately behind them, are known. Thismight provide an increased lift coefficient before stall is reached.Various vortex generators and fences are used to delay the onset of astall or to isolate the portion of a wing that is stalled. Furthermore,various stall warning/actuators allow aircraft to operate relativelyclose to their stall speed. Additionally, some combinations of airfoilsand wing configurations feature gentle stalls and so the vehicle can beoperated at the stall edge without abrupt drag increases or liftdecreases during the onset of a stall. Experimental aircraft have evenemployed rotary devices to permit low-speed flight, with mechanisms thatrestrict rotary moment and decrease drag or potentially augment liftwhen at higher speeds the wing provides the main lift. Many of the abovemechanisms provide this increased low-speed control at the expense ofweight and reliability.

In some high-tech aircraft, highly-active control is used to maintainstable operation over a wide range of speeds and orientations. Thisemulates the flying characteristics of natural fliers that change wingand airfoil geometry. In aircraft, such systems are complex, potentiallyheavy, and expensive, as well as fault-intolerant.

Airframe Configuration

The requirements for wide speed range, low power, light weight,unimpeded communications platform view, simplicity, and reliabilitypresent significant tradeoff challenges. A highly cambered airfoil helpswith lowering minimum flight speed, but is accompanied by a largenegative pitching moment that impacts the aeroelastic effects of wingtwist.

Furthermore, there is an inherent relationship between an aircraft'soverall airframe geometry and the design of its airfoils and controlsurfaces. Typical aircraft offset negative (i.e., nose-down) pitchingmoments through the use of tail moments (i.e., vertical forces generatedon the empennage with a moment arm being the distance from the wing tothe empennage) or through the use of a canard in front of the wing that,for pitch stability, operates at a higher lift coefficient than the wingand stalls earlier. Tails mounted in the up-flow of wingtip vortices canbe much smaller than tails positioned in the wing downwash, but thereare structural difficulties in positioning a tail in the up-flow.

Commercial airliners address the high coefficient of lift (C_(L))requirements for landing and takeoff with a complex array of slats andflaps that are retracted during high-speed flight to lower drag andgust-load severity. A rigid wing structure, and pitch controllabilityfrom the tail's area and moment arm, permit this approach. However, thisapproach is contrary to the requirement that the present aircraft carryfuel adequate to last for extended periods of time, and still beeconomical.

The very special requirements and technological challenges for theaircraft of the present invention have not been met by existing aircraftdesigns. Accordingly, there has existed a definite need for alightweight aircraft capable of both stationkeeping and flight over awide range of speeds, that consumes low levels of power for an extendedperiod of time, that supports a communications platform with a wide,unobstructed view, and that is characterized by simplicity andreliability. Embodiments of such an invention can serve as high altitudeplatforms. Embodiments of the present invention satisfy variouscombinations of these and other needs, and provide further relatedadvantages.

SUMMARY OF THE INVENTION

The present invention provides aircraft, aircraft components andaircraft subsystems, as well as related methods. Various embodiments ofthe invention can provide flight over a wide range of speeds, consuminglow levels of power for an extended period of time, and therebysupporting a communications platform with an unobstructeddownward-looking view, while and having simplicity and reliability.

In one variation, a wing of the invention is characterized by havingadequate camber to achieve a lift coefficient of approximately 1.5 atthe Reynolds number experienced by sailplanes or flexible-wingedstratospheric aircraft. The wing defines a leading edge and a trailingedge, and the trailing edge includes either a reflexed portion or atrailing edge flap that can extend upward. Either the reflexed portionor the flap is configured to provide the wing with a pitching momentgreater than or equal to zero in spite of the camber. This featureadvantageously allows for low-speed flight with a flexible wing in manyembodiments.

This feature is augmented by an extendable slat at the leading edge ofthe wing. These features, in combination, provide for an excellentcoefficient of lift of the wing, typically increasing it by more than0.3, and preferably by 0.4 or more, at airspeeds just above the stallspeed. Using its retractability, the slat can become part of the wing'sairfoil that is otherwise defined by the wing's camber. Slats areconvenient because they have a negligible or beneficial effect on awing's pitching moment. While flaps might help increase the C_(L) morethan slats, they do so at the cost of a big increase in negativepitching moment that potentially requires heavy, drag-producingcountermeasures for compensation.

In another variation of the invention, an aircraft comprises a flyingwing extending laterally between two ends and a center point,substantially without a fuselage or an empennage. The wing is swept andhas a relatively constant chord. The aircraft also includes a powermodule configured to provide power for the aircraft, and a supportstructure including a plurality of supports, where the supports form atetrahedron. This tetrahedron has corners in supportive contact with thewing at structurally stiff or reenforced points laterally intermediatethe center point and each end. The tetrahedron also has a corner insupportive contact with the wing's center point, which is alsostructurally stiff or reenforced. Advantageously, the flying wing isconfigured with a highly cambered airfoil and with reflex at a trailingedge. The wing is also configured with slats. These features providemany embodiments with the capability of high-altitude flight with a widerange of speeds.

A third variation of the invention is an aircraft, and its related powersystem, for generating power from a reactant such as hydrogen. The powersystem includes a fuel cell configured to generate power using a gaseousform of the reactant, the fuel cell being configured to operate at apower-generation rate requiring the gaseous reactant to be supplied atan operating-rate of flux. The power system also includes a tankconfigured for containing a liquid form of the reactant, wherein thetank includes a heat source for increasing a boiling-rate of thereactant. The tank is configured to supply its reactant to the fuel cellat a rate determined by the boiling-rate of the reactant, and the heatsource is configured to increase the boiling rate of the reactant to alevel adequate for supplying the resulting gaseous reactant to the fuelcell at the operating-rate of flux. An advantage of such an aircraft isthat it provides for a minimized system weight, volume and complexity,while not excessively sacrificing power generation.

In a fourth variation of the invention, the power system of the thirdvariation includes a tank that comprises an inner aluminum tank linerhaving an outer carbon layer, an outer aluminum tank liner having anouter carbon layer, and connectors extending between the inner and outeraluminum tank liners to maintain the aluminum tank liners' relativepositions with respect to each other. The volume between the inner andouter tank liners is evacuated to minimize heat transfer between thecontents of the tank and the outside environment. The connectors betweenthe inner and outer layers are configured with holes in their walls tominimize direct heat-conduction between the contents of the tank and theoutside environment.

In a fifth variation of the invention, an aircraft includes a hydrogensource, an oxygen source and a fuel cell configured to combine hydrogenfrom the hydrogen source and oxygen from the oxygen source to generatepower. The fuel cell is preferably configured to combine the hydrogenand the oxygen at less than one atmosphere of pressure, and morepreferably at roughly 2-3 psia. This advantageously allows stratosphericflight with simpler fuel cell technology.

Preferred embodiments of the above aspects of the invention, and variouscombinations of their features, provide for unmanned aircraft capable offlying in the stratosphere, in a stationkeeping mode, carrying a payloadof more than 100 kg that consumes more than 1 kw of electric power, andremaining aloft for a significant period of time while being able tooperate from a remote site where takeoff/landing weather is benign.

Other features and advantages of the invention will become apparent fromthe following detailed description of the preferred embodiments, takenin conjunction with the accompanying drawings, which illustrate, by wayof example, the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A is a perspective view of a first embodiment of an aircraftembodying features of the present invention, the aircraft having acowling removed to expose a fuel tank that the cowling conceals.

FIG. 1B is a front elevational view of the embodiment depicted in FIG.1A, having its cowling in place.

FIG. 1C is a right side elevational view of the embodiment depicted inFIG. 1B.

FIG. 1D is a to plan view of the embodiment depicted in FIG. 1B, rotatedby 90 degrees.

FIG. 2 is a system diagram of a fuel cell system for the embodimentdepicted in FIG. 1A.

FIG. 3 is a partial cross-sectional view of the fuel tank's wall in theembodiment depicted in FIG. 1A.

FIG. 4 is a partial cross-sectional view of a cross cell connector usedin the fuel tank's wall depicted in FIG. 3.

FIG. 5 is a cross-sectional view of a wing on the embodiment depicted inFIG. 1A.

FIG. 6 is a cross-sectional view of a wing on a variation of theembodiment depicted in FIG. 1A.

FIG. 7A is a top plan view of a third embodiment of an aircraftembodying features of the present invention.

FIG. 7B is a rear elevational view of the embodiment depicted in FIG.7A.

FIG. 8 is a top plan view of a fourth embodiment of an aircraftembodying features of the present invention.

FIG. 9 is a top plan view of a fifth embodiment of an aircraft embodyingfeatures of the present invention.

FIG. 10A is a top plan view of a sixth embodiment of an aircraftembodying features of the present invention.

FIG. 10B is a rear, elevational view of the embodiment depicted in FIG.10A.

FIG. 11A is a bottom plan view of a seventh embodiment of an aircraftembodying features of the present invention.

FIG. 11B is a front, elevational view of the embodiment depicted in FIG.11A.

FIG. 11C is a bottom plan view of a variation of the embodiment depictedin FIG. 11A.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS First PreferredAircraft Embodiment Fuel and Power Systems

A first preferred, high-performance aircraft embodiment 101, capable ofhigh-altitude stationkeeping within tight altitude and perimeterboundaries for extended periods of time, according to the presentinvention, is shown in FIGS. 1-4. The aircraft includes a wing 103, anempennage 105 and a plurality of motors 107. The empennage is preferablysuspended from the wing on an extension 109 to provide the moment armnecessary to control pitch and yaw. Thus, the extension's length will bebased on the empennage's surface area and the needed pitching and yawingmoments.

A fuel tank 111 is suspended below the wing using trusses and/or wires.A payload section 113 containing a communications payload 115 extendsforward from a lower portion 117 of the fuel tank, and is suspendedusing trusses, wires and/or supports 118. Preferably, the aircraftincludes a cowling or fuselage portion 119 (not shown in FIG. 1 toexpose contents) that forms a single aerodynamic body enclosing the fueltank and payload section.

Preferably, the wing 103 is unswept, extending 200 feet tip-to-tip. Thewing preferably has a constant 10 foot chord, and thus an aspect ratioof 20. The wing thus has an aspect ratio on the order of 20. Port andstarboard sides of the wing are each equipped with an inboard portion121 having no dihedral and an outboard portion 123 having a positivedihedral. The wing is torsionally flexible to limit the overall aircraftweight.

First Preferred Aircraft Embodiment Fuel and Power Systems

Each side of an inboard portion 121 of the wing mounts four electricmotors 107, and each side of an outboard portion 123 of the wing mountsfive electric motors, for a total of 18 electric motors. With referenceto FIG. 2, preferably the aircraft is powered by a hydrogen-air fuelcell system that uses gaseous hydrogen as fuel. The system includes afuel cell 131 that combines a reactant of gaseous hydrogen with oxygenand outputs electric power and water. The fuel cell powers an inverter133 that runs a motor 135 that drives a compressor 137 to compressoutside air to provide oxygen for the fuel cell. The air and hydrogencombine in the fuel cell to create the power both for the compressor'sinverter, and for an inverter 139 to run a propellor motor.

The fuel cell of the preferred embodiment can be configured to operatewith a fuel of gaseous hydrogen at reaction pressures (i.e., thepressure of the hydrogen gas when it reacts with the oxygen via amembrane electrode assembly) above 1 atmosphere, such as approximately15 psi or higher. However, unlike typical hydrogen-powered systems,which are designed with complex thermal and mechanical systems tooperate at reaction pressures of greater than one atmosphere, thepresent embodiment is preferably designed to operate at reactionpressures ranging below 1 atmosphere, and possibly down to even 2 or 3psia, when the aircraft is at cruise altitudes. This feature cansignificantly reduce the cost and weight of the power generation systemwhile increasing its reliability during high-altitude flight.

The hydrogen reacted by the fuel cell is preferably derived from a fuelsource comprising liquid hydrogen that is stored in the fuel tank 111.Storing the fuel as a liquid provides for the fuel to be stored in avolume that is small enough to fit reasonable aircraft shapes.Preferably, the cryogenic container(s) necessary to carry the fuel arerelatively lightweight. Other known hydrogen sources such as gaseoushydrogen tanks are within the scope of the invention.

As noted above, the fuel cell reacts air as an oxidizer, the air beingcompressed from ambient air. Preferably the air is taken into theaircraft at any given cruise flight condition so that storage andretrieval are not necessary. This allows for continuous operation (i.e.,operation for an unlimited period of time within the fuel capacity ofthe aircraft) of the aircraft at stratospheric flight conditions.

The oxidizer source for the aircraft preferably comprises an inlet forambient air and preferably a compression mechanism configured tocompress the ambient air. The compression mechanism is preferably acompressor as described above, but may also be other compressionmechanisms such as aerodynamic devices that operate using the rampressure generated by the aircraft's airspeed. Other known oxygensources such as oxygen tanks are also within the scope of the invention.

With reference to FIGS. 1A, 3 and 4, the fuel tank preferably includesan inner aluminum tank liner 151, having an inner carbon layer 153formed on it, and an outer aluminum tank liner 155, with an outer carbonlayer 157 formed on it. The internal radius of the inner aluminum layeris preferably four feet. Such a tank will preferably hold approximately1,180 pounds of liquid hydrogen.

Core cells 171 are bonded onto and extend between the inner and outeraluminum tank liners 151 and 155 to connect them. These cells arepreferably hexagonal, having vent holes 173 in the walls of the cells. Avacuum is created between the inner and outer aluminum tank liners,minimizing heat transfer between the fuel and the outside environment.The vent holes minimize the direct heat-conduction path. Preferably,each cell extends four inches between opposing sides. The fuel tankpreferably insulates the liquid hydrogen fuel so as to receive 28 orfewer watts through convection from the surrounding, ambient air.

The fuel cell is configured to operate at one or more power-generationrates that require the gaseous hydrogen to be supplied at relatedoperating-rates of flux. The heat received by the liquid hydrogen viaconvection through the insulated tank walls preferably causes the liquidhydrogen to boil at a boiling-rate lower than one or more (andpreferably all) of the anticipated boiling-rates desired to producegaseous hydrogen at the related operating-rates of flux. However, if ahybrid power system (e.g., a combination fuel cell and solar cellsystem) is used, there might be times when a zero boiling rate would bepreferred.

To provide hydrogen to the fuel cell at an acceptable rate over theconvection boiling rate, heat is either delivered to, or generated in,the fuel tank 111 by a heat source. That heat source is configured toincrease the boiling-rate of the liquid hydrogen to one or more desiredboiling-rates adequate to supply gaseous hydrogen to the fuel cell at anoperating-rate of flux. The fuel tank is configured to supply hydrogento the fuel cell at a rate related to and/or determined by theboiling-rate of the hydrogen, and thus operate the fuel cell at apower-generation rate adequate to power generation needs.

Preferably the heat source is an electrical heating element. The fuel inthe fuel tank is preferably boiled off over ten or more days to maintainthe aircraft's flight for at least that length of time. Preferably 1.5kilowatts of heater power are required to boil the liquid hydrogen offover that period of time. The heater is preferably configured such thatincreased levels of heater power are readily available when needed.

The aircraft is preferably includes a controller configured to regulatethe reaction pressure of either of (or both of) the fuel and theoxidizer. The controller regulates these reaction pressures in responseto the power requirements of the aircraft, increasing the givenpower-generation rate by appropriately increasing the reaction pressuresof one or both of the reactants. The controller preferably regulates thereaction pressure of the oxidizer by regulating the amount that thecompressor compresses the ambient air being fed to the fuel cell.Similarly, the controller preferably regulates the reaction pressure ofthe fuel being fed to the fuel cell by regulating the heat generated bythe heat source, and thus the boiling rate of the liquid fuel.

Under the regulation of the controller, at near sea-level conditions(e.g., conditions during take-off at normal ground-level airports), thecompressor might operate to compress the ambient air to reactionpressures above 1 atmosphere. However, compression will only occur if itis needed for adequate power generation. As the aircraft ascends tohigher altitudes, such as cruise-level altitudes, the ambient air islower in pressure. Rather than increasing compression levels tocompletely compensate for the lower ambient pressures, the controllerregulates the compression of ambient air to produce reaction pressuresbelow 1 atmosphere.

Under the regulation of the controller, at near sea-level conditionssuch as take-off conditions, the heat source will generally operate toproduce hydrogen gas at a reaction pressure that might exceed 1atmosphere. However, to maintain the fuel supply, hydrogen will only beprovided at a rate needed to generate the power required by the aircraftfrom the fuel cell.

During the flight of the aircraft, the controller preferably regulatesthe reaction pressures of the oxygen and the hydrogen such that hydrogenis reacted at the minimum rate necessary for the fuel cell to meet itspower-generation rate requirements. This minimizes both fuel use and thepower drain from boiling the liquid hydrogen. The controller alsopreferably regulates compressor operation to minimize power usage, whichtypically means that the air is compressed as little as possible.However, to protect the integrity of the fuel cell membrane, thereaction pressure of the fuel is maintained at no greater than apredetermined increment away from (and generally above) the reactionpressure of the oxidizer.

Preferably, at stratospheric conditions (e.g., 55,000-70,000 ft), whereambient air will typically be at less than 2 psia, the controller willtypically regulate the compressor to produce an air reaction pressure ofapproximately 6 psia (i.e., 6 psia, within a range established bysignificant digits). Preferably the increment of the fuel reactionpressure from the oxidizer reaction pressure will be no more thanapproximately 4 or 5 psi. Thus, at stratospheric conditions thecontroller will typically regulate the fuel source to produce a fuelreaction pressure of no more than approximately 10 or 11 psia.

Based on the recited fuel and propulsion system, it is estimated thatthe aircraft, with a gross weight of 4,000 pounds, can loiter at 60,000feet MSL within an area of 3,600 feet, with a speed of 130 feet persecond, and a potential dash speed of 180 feet per second whennecessary. To maintain a presence within the loiter diameter, theaircraft will bank up to 15 degrees in turning maneuvers.

First Preferred Aircraft Embodiment Airfoil Camber

With reference to FIG. 5, the wing of the preferred embodimentpreferably includes a highly cambered airfoil 201 that provides for highlift in low-speed flight regimes. The airfoil's camber permits theairfoil to achieve a lift coefficient of about 1.5 at the Reynoldsnumber typically experienced by sailplanes and the stratosphericaircraft of the invention.

An important aspect of the use of highly cambered airfoils is that theycause large negative pitching moments on the wing. In an alternatevariation, this embodiment uses flaps 203 to produce a highly camberedairfoil in low-speed flight regimes (see, FIG. 6).

First Preferred Aircraft Embodiment Variation One—Stiff Wing

In a first variation of the first preferred embodiment, the empennageprovides moments to react the overall moment of the airfoil's negativepitching moment. The supports 118 and the wing structure in the areaoutboard of the wing's connection to the slats provide structuralsupport and rigidity to the wing so as to avoid excessive wing torsion

First Preferred Aircraft Embodiment Variation Two—Counteracting Moments

In a second variation of the first preferred embodiment, the wingincludes slats and/or a reflexed trailing edge, providing positivepitching moments to react the overall effect of the airfoil's negativepitching moment.

In this variation, the wing of the preferred embodiment includes leadingedge slats 205 that extend conventionally. The slats increase the C_(L)before stall, and can optionally be deployed autonomously by relativewind directions and/or pressures. The deployment of some slat variationscan produce slight pitching moment change. While this effect is notrelevant for ordinary aircraft, it is important in the presentembodiment in preventing significant pitch-down wing torsion, such ascaused by the use of highly cambered airfoils (or flaps, as shown inFIG. 6). Furthermore, the slats let the airfoil of the invention achievea higher C_(L) in slower flight, and can retract at higher speeds to cutdrag and limit gust loads.

The use of a slat 205, preferably extending autonomously in high C_(L)flight regimes and retracting in low C_(L) flight regimes, can increasethe max C_(L) as much as 0.4 or even more, while having either anegligible or even a preferable effect on the pitching moment. Withcareful design and execution, the drag at both lower and higher speedscan be minimized, as demonstrated by airliners incorporating slattechnology.

Additionally, in this variation the wing's airfoil 201 incorporates areflexed portion 207 at its trailing edge, and is preferably configuredto produce a net zero or slightly positive pitching moment even thoughthe wing has high camber. In effect this simulates a standarddownward-loaded tail, with a very short moment arm, in the airfoilitself. Such airfoils can achieve a high maximum C_(L) for lower speeds,with reasonably low drag at higher speeds, while avoiding the wingtorsion problems caused by flaps.

Other Embodiments in General

Preferred embodiments of the invention have a variety of potential uses,the primary one being to carry a radio relay station that facilitatescommunications between ground, air, and/or satellite entities. For radiorelay purposes, such embodiments must support an antenna platform thatis horizontally (and azimuthally) stabilized and can “see” out in alldirections 25° below the horizontal without a wing or tail obstructingthe view. Optionally, the antenna platform can be lowered for use duringflight and raised to avoid contacting the ground during landings,takeoffs and taxiing.

A common role for embodiments of the aircraft will be to substitute forsolar-powered aircraft, such as the one disclosed in U.S. Pat. No.5,810,284 (the '284 patent), that cannot stationkeep for part or all ofthe year in some locations due to strong winds and/or limited solarradiation, such as is associated with long nights and low angles ofavailable sunlight during the winter at high latitudes.

The preferred aircraft is unmanned, and can stationkeep closely within alimited boundary. Being unmanned, the aircraft is preferably controlledeither by an autonomous system or by remote piloting.

In order to stationkeep closely, the aircraft will be slow flying whenwinds are light and it will generally maneuver continuously. Theaircraft also will fly sufficiently fast enough to stationkeep in strongwinds, and to fly significant distances to landing fields having benignweather conditions. It also will have enough climbing ability at peakaltitudes when fully loaded (i.e., at the early, fully fueled stage ofthe flight) to maintain its altitude in atmospheric down-currents.

The preferred aircraft will ordinarily stationkeep in the vicinity of analtitude between 55,000-70,000 feet. The available speed range willrange from a stall speed at less than 20 mph IAS (indicated airspeed) tomore than 40 mph IAS, which is about 70-140 mph TAS (true airspeed) at65,000 feet.

Other Embodiments in General Fuel and Power Systems

Preferred embodiments of the invention are fueled by liquid hydrogenreacted with atmospheric oxygen in a fuel cell. This fuel provides ahigh energy content. Thus, these embodiments preferably can operate evenin zero sun conditions. Other embodiments, including variations of thosedescribed above and below, can be configured to use other fuels, andpreferably to use gaseous fuels that are stored in liquid form.

Optionally, embodiments of the aircraft can include solar cells toprolong its flight in conditions having extensive available solarradiation. Furthermore, other hybrid combinations of power sources canbe used, including ones using regenerative fuel cells and/orconventionally combusted fuels (e.g., turbines or reciprocating engines)and they are within the scope of the invention. A conventionallycombusted fuel would preferably draw oxygen for combustion from thesurrounding air (usually with some compression).

The mechanical power generated by the power sources can directly driveeither a propeller or an attached generator that provides electricityfor propeller-driving electric motors. A generator may well be neededfor communication, control, and payload operation. A multiple-motorcontrol logic unit can mix power from multiple power sources as eachsituation requires. Additionally, embodiments will preferably have asmall battery energy system to provide redundant power for vehiclecommunication and control. This battery power can also be used to makelanding maneuvers safer.

Other Embodiments in General Configuration

In part, the invention pertains to the overall vehicle geometry. Theconfiguration of each embodiment is subject to numerous tradeoffconsiderations. The low-speed flight capability is preferablyaccommodated through the use of low aircraft weight, large wing area andhigh maximum lift coefficients of the wing airfoil. The power requiredat lower speeds is minimized by using a large wingspan that reducesinduced drag. High-speed flight is preferably accommodated through theuse of higher power generation rates, lower lift coefficients of thewing airfoil, smaller wing area, an extremely clean design and exteriorstructure, as well as appropriately designed propeller(s). The shiftingof the aircraft's CG (center of gravity) and the varying of theaircraft's rotational inertia as the fuel is consumed can be limited byappropriate fuel tank management.

Through the use of larger airfoil chords for a given span, larger wingareas and reduced stall speeds can be achieved. There is also a slightlydecreased power requirement, although the added weight of a “fat” wingmay negate these benefits. Nevertheless, for the preferred role ofembodiments of the present invention, a slower flight speed, even at thecost of extra power, can decrease the extent of the maneuveringnecessary to stationkeep during low wind speeds, and thereby increaseefficiency. Depending on the operational requirements, a normaloptimization study can determine the most useful chord compromise for afinal design.

To accommodate these conflicting design criteria, and thereby providefor a large speed range in preferred embodiments, the aircraft ispreferably characterized by a geometry change between the low-speed andhigh-speed flight regimes. The extent of the speed range will varydepending on the stationkeeping requirements. Less-stringentstationkeeping requirements (both laterally and vertically) can permitthe aircraft to operate with more efficient, gentle turns and to move todifferent altitudes if the wind profile showed a benefit in doing so,thus requiring less of a speed range than tighter stationkeepingrequirements.

Typically, a conventional vehicle is given pitch and yaw stabilityprimarily by a large tail moment (the tail forces times the moment armbetween the wing and the tail) and/or by a canard in front of the wingthat, for pitch stability, operates at a higher lift coefficient thanthe wing and stalls earlier. Tails mounted where they are in the up-flowof wing tip vortices can be much smaller than normal tails positioned inthe wing downwash, but there are structural difficulties with such“outboard tails.”

As a fuel load is consumed, the aircraft's CG (center of gravity) androtational inertia will vary. This effect can be limited by appropriatefuel tank management.

Other Embodiments in General Airframe Components

In part, the invention pertains to the specific design of aircraft'sairfoils. A torsionally flexible wing is characteristic of manyembodiments of the present invention. The typical airfoil of theinvention has enough camber to permit it to achieve a lift coefficientof about 1.5 at the Reynolds number typically experienced by theaircraft. As noted above, there preferably is some geometry change ofthe aircraft between the low-speed and high-speed flight regimes.Camber-changing devices are relatively simple and useful devices forchanging airfoil geometry.

An important aspect of the use of either flaps or highly camberedairfoils designed for high lift, is that such flaps (when extendeddownward) and airfoils cause a large negative pitching moment on a wing.This affects both the aircraft's overall stability and the wing'storsional deflection. Such wing twist at the outer portions of the wing,due to a negative pitching moment, can be a severe problem withtorsionally flexible, long-span wings such as are common in preferredembodiments of the invention.

Regarding the aircraft's overall stability, this problem can be handledby canard or tandem or tailed aircraft approaches within the scope ofthe invention. The configuration can produce enough pitch stability toovercome the negative pitch effect of the airfoil. The front surfaceneeds to have less percentage lift increase due to a small upward gustthan does the rear surface. This is accomplished by having the frontsurface operate at a higher C_(L) than does the rear. Note that the rearsurface is operating in the downwash wake of the front surface. For thestandard configuration this merely decreases the stabilizing effect ofthe tail, but the vehicle is still stable. For canard configurations thedownwash effect becomes much more troublesome, and dictates much higherC_(L)s for the front surface than for the rear, creating both overallvehicle inefficiencies and stall problems.

The larger problem caused by negative pitching moments is that, for atorsionally flexible wing the wing can twist significantly under thepitching moment. This twisting can even produce net negative lift in theouter wing, which is the cause of the undesirable aileron-reversaleffect.

Many embodiments of the present invention incorporate flexible wingdesign aspects such as those disclosed in the '284 patent. Various ofthese embodiments use one or more mechanisms to counteract this problem.

As noted in the first preferred embodiment, slats provide a mechanism tocounteract negative pitching moments in some embodiments, as well asincreasing the C_(L) (coefficient of lift) by 0.3, or even as much as0.4 or more before the onset of stall. Likewise, as used in the firstpreferred embodiment, a reflexed airfoil further counteracts thenegative pitching moment in some embodiments. With careful design, oneor both of these mechanisms can be used to achieve a high maximum C_(L)for lower speeds with reasonably low drag at higher speeds. Vortexgenerators can be used on the rear, underside of slats to inducevortices that may permit still higher maximum C_(L)s.

Other mechanisms are provided in some embodiments to limit the effectsof a negative pitching moment, as discussed in more detail in theadditional preferred embodiments below. These include “section” tails orcanards and swept flying wings.

The slatted, highly cambered and reflexed airfoil can be used in bothstandard-aircraft type embodiments of the invention, such as the firstpreferred embodiment, and also in flying wings. If the wing of a flyingwing is swept, it causes more pitch damping and stability. This alsomakes CG changes from fuel withdrawal from elongated-fore-aft tanks moretolerable.

Additional Preferred Embodiments

Both the first preferred embodiment and the additional preferredembodiments below are to be understood as including variationsincorporating different combinations of the power system and aircraftcomponent features described in this specification. Individual detailssuch as the number and placement of the motors are not depicted in someof the figures for simplicity.

Second Preferred Aircraft Embodiment

The second preferred embodiment of the invention incorporates variouscombinations of the above-described features into an aircraftincorporating the structural features of the span-loaded flying wingdisclosed and/or depicted in the '284 patent. Of particular note,variations of this embodiment incorporate a fuel cell operating atpressures described above with regard to the first preferred embodiment.Additionally, variations of this embodiment incorporate a fuel cellstorage tank configured to contain liquid hydrogen, and a heater to boilthe liquid hydrogen at a determined or predetermined boiling rate.

This aircraft is characterized by very flexible wing segments thattypically have a very slight positive pitching moment by virtue of theairfoils selected. While variations of the second preferred embodimentcan include highly cambered airfoils, flaps, slats and/or reflexedtrailing edges, this embodiment has not been found to be a highlyefficient platform for using high camber.

Third Preferred Aircraft Embodiment

With reference to FIGS. 7A and 7B, in this embodiment a wing 301 isdivided into a number of subsections 303, six being shown in the figure.Each subsection has a tail that permits the negative pitching moments ofthat section's highly cambered airfoil (or flap) to be reacted. The fouroutboard sections preferably have separate tails 305, and the twoinboard sections share a laterally extending tail 307. Optionally, thesectional structure of this preferred embodiment can adopt many of thefeatures and characteristics of the previous preferred embodiment and/orthe aircraft disclosed in the '284 patent.

In this multi-tail assembly, each of two symmetrically located “bodies”or fins 309 holds a liquid hydrogen storage and fuel cell system. Twosystems are preferably used for both symmetry and reliability. The twofins support the shared laterally extending tail 307. The two fins alsosupport landing gear, and a communications platform 311, which extendsdownward for better unobstructed viewing, can be retracted upward forlanding.

It should be noted that, as this embodiment rolls in flight, theoutboard subsections will tend to orient relative to the local flowregime and thus decrease the roll damping. Active control of the tailson the end subsection units can be used to eliminate the problem.However, the use of active control systems does increase the complexityof the system and thereby reduce its reliability.

The aircraft preferably has enough tails distributed across the wing 301to handle the pitching moment for each of the wing's subsections 303,providing for both vehicle pitch stability and limited wing twist. Ifthis embodiment's wing is designed torsionally stiff enough to keep thewing from significantly twisting under section pitching momentinfluences, then some or all of the four outer, separate tails 305 canbe removed and the central, laterally extending tail 307 can providevehicle pitch stability, even with flap deployment.

Fourth Preferred Aircraft Embodiment

With reference to FIG. 8, in this embodiment a conventional aircraftlayout is provided with a flexible wing 401, which supports a fuselage403 and is divided into a number of subsections 405. Similar to thethird preferred embodiment, each subsection has a wing-tail 407 thatpermits the negative pitching moments of that section's highly camberedairfoil (or flap) to be reacted. The main concern of the wing-tails isto prevent local wing torsion, as the overall aircraft pitching momentscan be reacted by a tail (not shown) mounted on the fuselage.

Because the wing is flexible, the roll damping is decreased as the wingtwists during roll. This effect can be decreased if the sections rotateon a strong spar, and both tip sections are rigidly attached to thetorsionally stiff spar to provide roll damping. The wing could be sweptin variations of this embodiment.

Fifth Preferred Aircraft Embodiment

With reference to FIG. 9, in this embodiment a long and typicallyflexible wing 501, such as might be found in the third preferredembodiment, is connected to a laterally extending tail 503 by aplurality of small “fuselages,” 505 some of which could simply be spars.The tail extends laterally across substantially the entire wing. Twoprimary fuselages 507 preferably include fuel and power modules. Theaircraft thus has pitch stability all across the span, even with the useof flaps on the wing.

The torsional flexibility of the wing and tail sections of thisembodiment will need to be made adequately rigid enough to limitdeflection during roll unless active control is to be used. As notedabove, it is preferable to avoid active control if possible.

As previously noted for all embodiments, this embodiment can includevariations having different combinations of slats and flaps (e.g.,slotted flaps). These include variations characterized by the tailhaving a small chord and zero lift at intermediate speeds.

Sixth Preferred Aircraft Embodiment

With reference to FIGS. 10A and 10B, in this embodiment a long andtypically flexible wing 601, such as might be found in the thirdpreferred embodiment, is connected to a laterally extending tail 603 bya plurality (namely four) of “fuselages,” 605, each being a fuel/powermodule that also provides an adequate moment arm to support the tail.Each outboard end 607 of the wing extends roughly 25 feet beyond theoutermost fuselage and is made torsionally strong enough such thatflap/aileron deflection is limited to about half of that used in theinner, span-loaded 90 feet of the wing. This construction provides thataileron reversal will only occur at speeds significantly higher thanpreferred indicated airspeeds.

Additionally, the four fuselages 605 provide mountings for simplelanding gear (e.g., two tiny retractable wheels on each fuselage). Aradio relay pod can be lowered during flight to a level where 30° banksof the aircraft will not obstruct the pod's visibility at more than 20°below the horizon.

Variations of the Third through Sixth Embodiments

Another approach within the scope of the invention is to vary theabove-described third through sixth embodiments to have canards ratherthan tails. It should be noted that a lower C_(L) is required on therear wing surface (i.e., it has an early stalling front surface). Thiswill likely cause higher levels of drag than the described variationswith tails.

Seventh Preferred Aircraft Embodiment

With reference to FIGS. 11A and 11B, a seventh, preferred embodiment isa swept flying wing design having a wing 701 and a 6-element tetrahedronframe 703 formed of compression struts. A fuel and power module 705 anda radio platform 707 are centrally located and preferably supported bythe tetrahedron frame. The tetrahedron frame adds great strength to theinner portions of the wing, permitting the weight of fuel and powermodule and the radio platform to be handled readily. Stabilizers and/orcontrol surfaces can optionally be mounted on the fuel and power moduleto add further stability and/or control.

Three elements of the tetrahedron frame 703 are preferably in a planedefined by the wing's main spars, extending along both sides of the wing701. Two wing-based elements 721 of these three spar-plane elementseither extend from a common, structurally reenforced point at the nose,along the spars, or are composed of the spars themselves. The third ofthe three spar-plane elements is a laterally extending element 723 thatextends between the spars from spar locations roughly 50 or 60 feetapart. There can be a benefit in integrating the lateral element of thetetrahedron into an extended wing chord in the middle portion of theaircraft (not shown). If this embodiment's span is 140 feet (having anaspect ratio approximately in the range of 14-17.5), the cantileveredwing elements outboard of the tetrahedron will laterally extend 40 or 45feet each, being a somewhat longer distance when considering the sweep,but still a relatively short distance that is consistent with goodtorsional and bending strength.

The remaining three elements of the tetrahedron frame 703 extenddownward to a common point 725. Two side-descending elements 727 ofthese three downward-extending elements extend down from the two ends ofthe laterally extending element 723, while the third, acenter-descending element 729, of these three downward-extendingelements extends down from the common, structurally stiff or reenforcedpart of the spars at the nose of the aircraft, where the two wing-basedelements 721 meet.

The drag of externally exposed compression struts is of aerodynamicrelevance, and these should be design aerodynamically. Omitting theportions of the compression struts that are within the wing, theremaining, exposed elements represent roughly 100 feet or less ofexposed strut length. With 1 foot chord, and a low drag shape giving aCd_(o) of approximately 0.01, only 1 ft² of equivalent flat plate areais added to the plane by the exposed elements.

In the relatively simple configuration of this embodiment, pitch and yawcontrol can be achieved by tip elevons, or more preferably, by wingtipsthat rotate about an axis along the wing's quarter chord. Thisrotating-tip type of control has been successfully implemented in flyingwings and conventional aircraft.

A benefit of many variations of the swept flying wing is that, byappropriate wing twist (and hence lift distribution) the tips can be ina region featuring upwash, letting the tips produce thrust andpermitting banked turns without causing adverse yaw. This isaccomplished without the drag of a vertical surface.

Furthermore, many variations of this embodiment will have strong pitchstability, thus providing the ability to accommodate a reasonablenegative pitching moment, such as from positive flaps that increasecamber. If these portions are forward of the CG, the configuration pitchstability is more readily able to accommodate the effects of airfoilpitch instability. Preferably this embodiment of the aircraft includes acambered airfoil with reflex and slats, taking full advantage of thestrong tetrahedron structure for distributing loads. Thus, thecombination of the cambered/reflexed/slatted airfoil, used on the flyingwing of the present embodiment, is especially preferred

With reference to FIG. 11C, in a variation of the seventh, preferredembodiment, two power pods 751 are located far out at the ends of thelaterally extending tetrahedron element 723, making the aircraft into aspan-loaded, swept flying wing.

A Further Variation of the Embodiments

The above described embodiments can each be varied so as to be directedto a multi-wing aircraft such as a biplane, such as with each winghaving half the chord of the equivalent monoplane wing. The vehicleperformance would remain about the same but the wing's negative pitchingmoment effect would be reduced because the chord would be halved. Thebig box truss has merit for achieving torsional and bending rigidity andperhaps for lower wing weight. Nevertheless, there is a drag penalty dueto the struts and wires and their intersections with the wing.

If a 100-foot span wing with an 8′ chord and thus a 12.5 aspect ratio(800 ft², at a high-speed C_(L) of 0.3 having a parasite dragcoefficient of 0.007 and hence a drag area of 5.6 ft²) were equated to abiplane with two 4-foot chord wings, having 600 ft of 1/16″ piano wireto stabilize the box formed by the two wings, the wire drag area wouldbe more than 3 ft². Considering strut drag, and the fact that the lowerReynolds number for the airfoils adds to their drag, the wing drag areawould more than double and inhibit high-speed flight for thatembodiment.

From the foregoing description, it will be appreciated that the presentinvention provides a number of embodiments of a lightweight aircraftcapable of both stationkeeping and flight over a wide range of speeds,while consuming low levels of power, for an extended period of time,while supporting an unobstructed communications platform, and whileexhibiting simplicity and reliability

While a particular form of the invention has been illustrated anddescribed, it will be apparent that additional variations andmodifications can be made without departing from the spirit and scope ofthe invention. Thus, although the invention has been described in detailwith reference only to the preferred embodiments, those having ordinaryskill in the art will appreciate that various modifications can be madewithout departing from the invention. Accordingly, the invention is notintended to be limited, and is defined with reference to the followingclaims.

1. An aircraft, comprising: a fuel source configured to provide a fuel;an oxidizer source configured to provide an oxidizer; a fuel cellconfigured to react the fuel with the oxidizer to operate at a givenpower-generation rate; and a controller configured to regulate thereaction pressure of a reactant of the group consisting of the fuel andthe oxidizer; wherein the controller is configured to regulate thereaction pressure to be less than one atmosphere with the aircraft at acruise altitude and the fuel cell operating at the givenpower-generation rate.
 2. The aircraft of claim 1, wherein the cruisealtitude is in the range of 55,000 to 70,000 feet.
 3. The aircraft ofclaim 1, wherein the reaction pressure of the reactant is not greaterthan 11 psia.
 4. The aircraft of claim 1, wherein the reaction pressureof the fuel is not greater than 11 psia, and wherein the reactionpressure of the oxidizer is not greater than 11 psia.
 5. The aircraft ofclaim 1, wherein the reaction pressure of the reactant is not greaterthan 10 psia.
 6. The aircraft of claim 1, wherein the reaction pressureof the fuel is not greater than 10 psia, and wherein the reactionpressure of the oxidizer is not greater than 10 psia.
 7. The aircraft ofclaim 1, wherein the reaction pressure of the reactant is not greaterthan 6 psia.
 8. The aircraft of claim 1, wherein the reaction pressureof the reactant is approximately 6 psia and the cruise altitude is inthe range of 55,000 to 70,000 feet.
 9. The aircraft of claim 1, whereinthe controller is further configured to regulate the reaction pressureof the reactant in response to the power requirements of the aircraft.10. The aircraft of claim 1, wherein the oxidizer source comprises aninlet for ambient air and a compression mechanism configured to compressthe ambient air.
 11. The aircraft of claim 10, wherein the controller isfurther configured to regulate the reaction pressure of the oxidizer byregulating the amount by which the compression mechanism compresses theambient air.
 12. The aircraft of claim 1, wherein the fuel sourcecomprises a hydrogen tank containing liquid hydrogen, and a heat sourcefor controllably boiling the liquid hydrogen.
 13. The aircraft of claim12, wherein the controller is further configured to regulate thereaction pressure of the fuel by regulating the rate at which the heaterboils the liquid hydrogen.
 14. The aircraft of claim 1, wherein thecontroller is further configured to regulate the reaction pressure ofthe fuel to be no greater than a predetermined increment above thereaction pressure of the oxidizer.
 15. The aircraft of claim 1, wherein:the oxidizer source comprises an inlet for ambient air and a compressionmechanism configured to compress the ambient air; the fuel sourcecomprises a hydrogen tank containing liquid hydrogen, and a heat sourcefor controllably boiling the liquid hydrogen; and the controller isfurther configured to regulate the reaction pressure of the oxidizer byregulating the amount by which the compression mechanism compresses theambient air, and to regulate the reaction pressure of the fuel byregulating the rate at which the heater boils the liquid hydrogen; andthe controller is further configured to regulate the reaction pressuresof the fuel and the oxidizer such that the power-generation rate of thefuel cell varies in response to the power requirements of the aircraft,and the reaction pressure of the fuel is no greater than a predeterminedincrement above the reaction pressure of the oxidizer.
 16. The aircraftof claim 15, wherein, with a cruise altitude in the range of 55,000 to70,000 feet, the reaction pressure of the oxidizer is approximately 6psia, and the predetermined increment is approximately 4-5 psi.
 17. Anaircraft, comprising: a hydrogen source including a hydrogen tank and amechanism configured to regulate delivery of hydrogen from the hydrogentank; an oxygen source including a compression mechanism configured tocompress ambient air from outside of the aircraft; a fuel cellconfigured to react hydrogen from the hydrogen tank with oxygen from thecompression mechanism to generate power; and a control system configuredto control the operation of the hydrogen source and the oxygen source ata given aircraft flight condition such that the fuel cell reacts oxygenat a first reaction pressure with hydrogen at a second reactionpressure, wherein the first reaction pressure is less than oneatmosphere, and wherein the difference between the first reactionpressure and the second reaction pressure is no greater than apredetermined limit.
 18. The aircraft of claim 17, wherein the secondreaction pressure is less than one atmosphere.
 19. The aircraft of claim17, wherein the control system is configured to vary the first andsecond pressures based on power requirements of the aircraft.
 20. Theaircraft of claim 19, wherein the control system is configured such thatat a stratospheric flight condition, the first pressure is approximately6 psia, and the predetermined limit is not greater than 5 psi.